The present disclosure relates to a gas turbine engine and, more particularly, to a combustor section therefor.
Gas turbine engines, such as those that power modern commercial and military aircraft, generally include a compressor section to pressurize an airflow, a combustor section to burn a hydrocarbon fuel in the presence of the pressurized air, and a turbine section to extract energy from the resultant combustion gases. Relatively high temperatures are observed in the combustor section such that cooling airflow is typically provided to meet desired service life requirements.
The combustor section typically includes a combustion chamber formed by an inner and outer wall assembly. Each wall assembly may include a support shell lined with heat shields often referred to as liner panels. In some combustor chamber designs, the combustor includes liner panels with a hot side exposed to the gas path and an opposite, or cold side, that has features such as cast in threaded studs to mount the liner panel and a full perimeter rail that contact the inner surface of the combustor liner support shells. The wall assemblies are segmented to accommodate growth of the panels in operation and for other considerations. Combustor panels typically have a quadrilateral projection (i.e. rectangular or trapezoid) when viewed from the hot surface. The panels have a straight edge that forms the front or upstream edge of the panel and a second straight edge that forms the back or downstream edge of the combustor. The panels also have side edges that are linear in profile.
The liner panels extend over an arc in a conical or cylindrical fashion in a plane and terminate in regions where the combustor geometry transitions, diverges, or converges. This may contribute to durability and flow path concerns where forward and aft panels merge or form interfaces. These areas can be prone to steps between panels, dead regions, cooling challenges and adverse local aerodynamics.